It is known that aircraft, and in particular jumbo jet civil aircraft, are equipped with automatic flight controls comprising a flight control computer (FCC) which groups the automatic pilot and the flight director. These automatic flight controls comprise, inter alia, a speed reference system, particularly active in take-off and go-around phase and delivering piloting information to the flight director and/or to the automatic pilot.
In addition, it is known that the flight director comprises an attitude director indicator controlled by said speed reference system and whose purpose is to display the attitudes of the aircraft; to that end, said attitude director indicator comprises a model and a mobile sphere, the position of said model with respect to said mobile sphere materializing the real attitude information of the aircraft, whilst the information of the flight director, furnished by the FCC computer, is displayed by horizontal and vertical bars, including a pitch bar, which indicates the tendency of the longitudinal control of the aircraft. The position of this pitch bar with respect to the model indicates to the pilot either an order to dive or an order to pull up, or, if said pitch bar is superposed on said model, that the aircraft lies in the desired configuration. In this way, the action of the pilot (or of the automatic pilot) is to act on the elevator, in order to superpose the pitch bar on the model.
It will be readily appreciated that, in the event of failure of an engine, such a speed reference system considerably simplifies the pilot's work which in that case consists in acting on the elevator to maintain the pitch bar and the model of the attitude director indicator in superposition.
Modern aircraft are thus equipped with such a speed reference system. For example, the twin-engine aircraft AIRBUS A-310 and A-300-600 comprise a speed reference system essentially constituted by a voter (i.e. a comparator with three inputs, at the output of which appears that of the three input signals whose amplitude is included between those of the other two), which receives at its inputs electrical signals respectively representative of a first difference between a desired speed displayed by the pilot and the real aerodynamic speed of the aircraft (corrected by the dynamic pressure), of a second difference between the real pitch attitude and a reference pitch attitude (for example selected to be equal to 18.degree.), and of a third difference between the real baroinertial vertical speed and a vertical speed limit (for example determined as being the minimum climbing speed with one engine at the gradient of 2.4%).
A vote between these three differences is made at each step of calculation of the FCC flight control computer.
As will be seen hereinafter in greater detail, during a vote:
(a) if the flight conditions upon take-off are normal, i.e. if the thrust/mass ratio concerning the aircraft is high, with the result that the aircraft continuously increases its kinetic energy and its potential energy, the first difference is negative, whilst the third is positive, with the result that the second difference is close to zero. The aircraft is then controlled to maintain an attitude equal to the reference attitude, with an increasing speed greater than the desired speed displayed by the pilot; PA1 (b) if the thrust/mass ratio is low, either because the mass of the aircraft is high, or because an engine has broken down, the first difference becomes zero, whilst the second becomes negative and the third remains positive, with the result that it is this first difference which is transmitted by the voter. Under these conditions, the action of the pilot or of the automatic pilot is therefore to actuate the elevator of the aircraft to modify the attitude thereof in order to maintain the aerodynamic speed of the aircraft equal to the desired speed displayed by the pilot, increased by 10 kts, the attitude then being less than the reference attitude. In that case, the potential energy of the aircraft increases with constant kinetic energy.
A piloting system of the type which has been briefly described hereinabove is essentially designed to take into account the cases of failure of an engine. It goes without saying, since it uses real parameters of the aircraft (corrected aerodynamic speed, attitude and baro-inertial vertical speed) which are sensitive to the aerodynamic environment thereof, that this piloting system also takes into account the disturbances of this environment. However, this latter consideration is accessory and may not be fine enough for a survival piloting in the event of the aircraft encountering a minitornado.
"Minitornado" is understood to mean the meteorological disturbances commonly designated in aeronautics by the English words "windshear", "downburst" or "microburst", and will be referred to hereinafter as "windshear".
Such windshear is essentially constituted by violent eddying, descending, air streams whose speed may be greater than 10 m/s and which present considerable horizontal and vertical speed components.
Although the probability of the aircraft encountering such windshear during take-off or landing is low, nonetheless there is a danger that the flight of an aircraft be strongly disturbed by windshear during these flight phases, during which its safety margin is relatively sensitive. It is estimated that, in the last twenty years, windshear has been responsible for about thirty accidents or failures upon take-off and landing, involving more than 600 deaths.
Consequently, it is an object of the present invention to improve the speed reference system described hereinabove in order to render it even more sensitive to the possible aerodynamic disturbances of the environment of the aircraft, in order in particular to define a strategy of survival piloting in the event of windshear.